Hall Thrusters are typically used in rockets, satellites, spacecraft, and the like. In a typical Hall Thruster, the working fluid is plasma and the means of acceleration is an electric field. A conventional Hall thruster typically includes a plasma accelerator that includes a gas distributor and an anode located at one end of a channel. A gaseous propellant, a propellant that is gaseous at standard temperature and pressure, e.g., xenon, is introduced to the plasma accelerator. An electric circuit provides an electric potential applied between the anode and a floating externally located cathode which emits electrons. A magnetic circuit structure establishes a transverse magnetic field in the plasma accelerator presenting an impedance to electrons attracted to the anode. As a result, the electrons spend most of their time drifting azimuthally (orthogonally) due to the transverse magnetic field. This allows the electrons time to collide with and ionize the neutral atoms of the gaseous propellant. The collisions create positively charged ions accelerated by the electric field to create a flux of ions that may be used as an ion source or to create thrust. See e.g., U.S. Pat. Nos. 6,075,321, 6,150,764, and 6,735,935, 6,834,492 and co-pending applications U.S. patent application Ser. No. 10/761,565 entitled Multi-Functional Power Supply for A Hall Thruster, filed on Jan. 21, 2004, U.S. patent application Ser. No. 11/301,857 entitled A Hall Thruster With Shared Magnetic Structure, filed on Dec. 13, 2005, and U.S. patent application Ser. No. 11/412,619 entitled Combined Radio Frequency and Hall Effect Ion Source and Plasma Accelerator System filed on Apr. 27, 2006, by one or more common inventors hereof and all of the same assignee, all of which incorporated in their entirety this reference herein.
Condensable propellants, e.g., Bismuth, exist in a non-gaseous state at standard temperature and pressure. Solid condensable propellants must be heated to achieve a liquid state and then vaporized to produce the gaseous state necessary for use with a Hall thruster, which requires considerable energy. Using condensable propellants, however, can be advantageous over gaseous propellant because they are easy to collect and, in certain cases, non-toxic, inexpensive, and plentiful. Certain condensable propellants also have a higher atomic mass and larger electron impact ionization cross-section than common gaseous propellants, such as xenon. This results in more complete propellant utilization, higher thrust and improved specific impulse. When used with larger Hall thrusters with available energy, the benefits of using condensable propellants often outweigh the energy requirements of heating and converting the condensable propellants to a gaseous state.
One conventional Hall thruster that utilizes condensable propellants is disclosed in U.S. Pat. No. 7,059,111 to King, incorporated by reference herein. The thruster as disclosed by King controls the flow rate and vaporization rate of the condensable propellant using a gravity-fed method that delivers liquid condensable propellant to the vaporizer. The flow rate and vaporization rate of the condensable propellant are determined by controlling the temperature of the vaporizer. During operation, the vaporization process causes the temperature of the vaporizer to increase which increases vaporization and the density of the plasma discharge. This increases feedback of discharge current in the vaporizer, further increasing the heat of the vaporizer which, in turn, increases the flow rate and vaporization rate. Without some type of control, a run-away condition of vaporization rate and flow rate of the condensable propellant results. The King thruster relies on biasing a second anode to control the feedback of discharge current and regulate the temperature of the vaporizer and the resulting vaporization rate and the flow rate. Such techniques are complex and cumbersome, often resulting in the first or second anode melting, and typically run only at one operating condition. The King thruster also requires pre-heating of the anode and discharge chamber through a secondary gas discharge, e.g. xenon.
Other conventional Hall thrusters that utilize condensable propellants employ electromagnetic pumps utilized to control flow rate and vaporization rate, e.g., the Soviet TsNIIMASH TAL thruster as modified for the VHITAL program. Relying on electromagnetic pumps to control flow rate and vaporization rates is complicated and expensive.
Older conventional Hall thrusters that utilize condensable propellants employ external crucibles or vaporization chambers to control the vaporization rate external to the thruster, relying upon gravity to maintain the liquid vapor interface. Relying on gravity to maintain the interface is not a valid approach for space applications. Furthermore, vaporization external to the thruster results in large power losses due to radiation from the propellant flow lines. These losses reduce the system level efficiency.
Hall thrusters or ion sources which can utilize both condensable propellants, such as Bismuth and gaseous propellants, such as Xenon are known. See e.g., the King patent discussed above. However, to date, conventional Hall thrusters that utilize both condensable propellants and gaseous propellants do not store the condensable propellant and the gaseous propellant in a single storage vessel and are unable to control the flow rate of the condensable propellant and the gaseous propellant, or a mixture of the condensable propellant and the gaseous propellant.
A Hall thruster that uses a solid condensable propellant wherein the flow rate and vaporization rate are regulated by controlling the linear feed rate of the solid condensable propellant would be advantageous. To date, no such Hall thruster exists.